1. Field of the Invention
The present invention relates generally to turbine blades, and more specifically to an air cooled turbine blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially an industrial gas turbine engine, a compressor provides compressed air into a combustor to be burned with a fuel and produce a hot gas flow that is passed into a turbine to drive the compressor and, in the case of the IGT, to drive an electric generator. The efficiency of the engine can be increased by passing a higher temperature flow into the turbine. However, the turbine inlet temperature is limited to the material properties of the first stage airfoils (rotor blades and stator vanes) and to the amount of cooling provided for the airfoils.
In order to allow for higher turbine inlet temperatures—and, also to increase the engine efficiency—complex cooling circuits have been proposed for the vanes and blades. A combination of convection cooling, impingement cooling and film cooling is used to maximize the cooling effectiveness while minimizing the amount of compressed cooling air bled off from the compressor. Turbine airfoil designers are always trying to maximize the airfoil cooling capability while minimizing the amount of cooling air used.
FIG. 1 shows a prior art turbine blade with an internal cooling circuit that includes a leading edge cooling channel 11 with trip strips along the inner channel walls, and a blade tip cooling hole 15 at the end of the channel. The blade cooling circuit also includes a aft flowing triple pass or 3-pass serpentine flow cooling circuit (all convective cooling) with a first leg 21 connected to the cooling air supply that flows upward toward the tip, a second leg 22 that flows downward toward the root, and a third leg 23 that flows upward and positioned along the trailing edge region of the blade. A row of cooling air exit holes 24 are placed along the trailing edge and connected to the third leg of the serpentine flow circuit. Trip strips are also located along the serpentine passages. FIG. 2 shows a cross section view of the blade of FIG. 1 taken through the section A-A.
In the prior art blade of FIG. 1, compressed cooling air is supplied to the leading edge cooling channel 11 and the first leg 21 of the serpentine flow circuit from the pressurized cooling air source such as the compressor. The cooling air in the leading edge channel 11 flows upward toward the blade tip. The leading edge channel has a rough triangular cross section shape as seen in FIG. 2. The inner surface area of the leading edge cooling channel 11 is reduced in cross sectional area to an apex of an acute angle. As such, the distribution of the cooling flow to the leading edge corner decreases and the flow velocity as well as the heat transfer coefficient is comparatively reduced. The spent cooling air is then discharged at the blade tip section through the tip cooling opening 15 at the end of the channel as represented by the arrow in FIG. 1. Since the cooling channel has to satisfy the minimum ceramic core size in order for the blade to be cast, frequently the tip cooling hole ends up as an oversized cooling flow passage for the leading edge channel at the exit open region. In addition, the discharge open region at the blade tip location is also subject to the mainstream pressure variations. Mal-distribution of the cooling flow as well as mal-metal temperature at the blade leading edge upper channel location is evidenced in the engine hardware. Also, a single pass radial channel cooling is not the best way of utilizing the total amount of cooling air. In the FIG. 1 prior art blade cooling circuit, about 25% of the cooling air supplied to the blade flows into the leading edge cooling channel 11 while the remaining 75% flows through the serpentine flow circuit by entering the first leg 21. All of the cooling air flow through the leading edge channel 11 is discharged out the blade tip opening 15 and wasted.
FIG. 5 shows another prior art turbine blade with a similar internal cooling circuit to that of FIG. 1 having a separated leading edge cooling feed channel 11 for the airfoil that also feeds a leading edge backside impingement cavity 13 through metering and impingement holes 12, and showerhead cooling circuit 14. The blade in FIG. 5 also includes a blade tip discharge cooling hole 25 at the end of the third leg 23 in the 3-pass serpentine circuit. FIG. 6 is a cross section view of the blade of FIG. 5 taken along the line D-D in order to maintain the leading edge feed channel 11 through flow velocity high enough to overcome the rotational effect at the blade upper span, a considerable amount of cooling air has to be used in the through flow channel while cooling air is bled off from the through flow channel to provide the blade leading edge backside impingement and showerhead film cooling. This additional upper span cooling air is then discharged through the core print out holes 15 and 25 at the blade tip section as represented by the two arrows in FIG. 5 like in the FIG. 1 blade, the cooling air passing through the leading edge channel 11 that is not bled off through the showerhead film holes 14 is discharged out through the tip opening 15 and wasted.